1. Field of the Invention
The present invention generally relates to determining spacecraft attitude and, more particularly, to determining attitude using horizon-crossing indicators.
2. Description of the Related Art
In general, a spacecraft orbiting the earth randomly spins about a spin axis during its stay in the orbit. An attitude determination process can determine such spinning spacecraft""s orientation relative to a coordinate system associated with the earth or another reference body such as the sun. Generally, the process is carried out using an onboard sensor, such as a Horizon-Crossing Indicator (HCI) mounted on the spacecraft.
Horizon-crossing indicators are light-detecting sensors which are configured to sense the edge, or the limb, of the earth, which appears as an illuminated disc. Each time the sensor detects the change from light to dark, or dark to light, it produces a signal that is used as an indication of horizon crossing for attitude control and guidance of the satellite. In a current design, as illustrated in FIG. 1, an HCI sensor 10 has a cone shaped field of view 12 that is often referred to as the bore-sight of the sensor. The bore-sight 12 is centered about the bore-sight axis 14 of the sensor, i.e., focal axis of the HCI sensor. As the spacecraft 16 constantly spins about a spin axis 18, this cone-shaped bore-sight 12 sweeps the space. When the bore-sight of the HCI sensor 10 scans from the space on to the illuminated disc 20 of the earth 22 or vice versa, it senses the change in the light intensity and produces a signal. In this respect, the signal may be produced as the bore-sight 12 crosses the earth horizons at a first position 24 or earth-in position and also as the bore-sight crosses the earth horizons at a second position 26 or earth-out position.
If the spacecraft spins in equilibrium, the HCI sensor 10 scans a linear path between the earth-in and earth-out positions 24 and 26, and forms a so-called chord 28 of the illuminated disc 20. The chord width between the earth in and earth out positions 24 and 26, together with the spin period of the spacecraft and the known size of the earth 22, indicate how far above or below the center of the earth the HCI sensor is scanning. This information allows computation of the angle between the spin axis and another axis from the spacecraft to the center of the earth. This measurement, together with an observation of the sun, can be used to determine the attitude of the satellite 16.
As mentioned above, in the current state of the art, spinning satellite attitude determination with horizon-crossing indicators requires constant (i.e. equilibrium) spin about a single axis from which earth-chord widths are calculated. With the current state of the art, however, it is not possible to make attitude-determination for an arbitrarily spinning spacecraft (one that is not in equilibrium). If the spacecraft spins arbitrarily, the HCI sensor scans a non-linear path resulting in a non-linear chord width that makes it difficult to determine how far the scanning action occurs from the center of the earth. This, in turn, results in ambiguities in determining the attitude of the subject spacecraft.
In light of the foregoing, a need therefore exists in the current spacecraft technology for a method that permits unambiguous determination of the attitude of a spacecraft.
The present invention provides a method for determining the attitude measurement of a spacecraft that spins arbitrarily while orbiting the earth. In particular, and regardless of spin-axis motion of the spacecraft, the present invention uses an integrated angular-rate data, e.g., data in the form of a propagated quaternion in conjunction with the horizon crossing indicator HCI data from at least three time-tagged horizon-crossing measurements that determine the earth center.
In one aspect of the present invention, a process for determining an attitude measurement of a spacecraft using a sensor which senses the illuminated circle of a celestial body, comprises the steps of generating a first signal indicative of a first horizon-crossing time; generating a second signal indicative of a second horizon-crossing time; generating a third signal indicative of a third horizon-crossing time; and determining the center of the celestial, from which an estimate of the spacecraft attitude with an additional measurement of a second celestial body can be made.
These and other features, aspects and advantages of the present invention, will become better understood with reference to the following drawings, description and claims.